(1) Field of the Invention
The present invention relates to an improved design for a turbine blade to be used in a gas turbine engine.
(2) Prior Art
Referring now to FIG. 1, turbine blades 10 typically used in gas turbine engines include a platform 12, an airfoil 14 extending radially from a first side of the platform, and an attachment or root portion 16 extending from a second side or underside of the platform. The root portion 16 typically includes a dovetail portion with a plurality of serrations and a neck portion between the dovetail portion and the underside of the platform. As shown in FIG. 1, the airfoil 14 may overhang the footprint of the root portion 16. Also formed in the turbine blade 10 is a pocket structure 18, which is typically a cast structure. The neck portion of the attachment or root portion 16 begins just beneath the pocket structure 18 and forms a limiting structure in the sense that significant stresses act in this region—stresses which if not dealt with properly could be the source of cracks and other potential failure modes. Balancing stress concentrations between suction and pressure sides of the neck portion and the stress on the turbine airfoil 14 is highly desirable.
Given the lower speeds and temperatures of low pressure turbine airfoils, the root axial length of the root portion 16 is generally shorter than the airfoil chord axial component. Most low pressure turbine airfoils also have shorter attachment root neck lengths. The overhung airfoil and short neck length create a load path that will concentrate stress in the root in most cases. This is exemplified in FIG. 2. In certain cases, these stresses are unacceptable and a potential source of cracks. The traditional solution to this problem is to increase root axial length, width, and enlarge serration sizes. This traditional solution requires a new disk design and increases weight.